This invention relates to gas turbines and, more particularly, to a concept for efficiently cooling ultrahigh temperature turbine rotor blades.
It is well understood that gas turbine engine shaft horsepower and specific fuel consumption, which is the rate of fuel consumption per unit of power output, can be improved by increasing turbine inlet temperatures. However, current turbines are limited in inlet temperature by the physical properties of their materials. To permit turbines to operate at gas stream temperatures which are higher than the materials can normally tolerate, considerable effort has been devoted to the development of sophisticated methods of turbine cooling. In early turbine designs, cooling of high temperature components was limited to transferring heat to lower temperature parts by conduction, and air cooling was limited to passing relatively cool air across the face of the turbine rotor disks.
In order to take advantage of the potential performance improvements associated with even higher turbine inlet temperatures, modern turbine cooling technology utilizes hollow turbine nozzle vanes and blades to permit operation at inlet gas temperatures in the 2000.degree. to 2300.degree. F. (1094.degree. to 1260.degree. C.) range. Various techniques have been devised to air cool these hollow blades and vanes. These incorporate three basic forms of air cooling, either singly or in combination, depending on the level of gas temperatures encountered and the degree of sophistication permissible. These basic forms of air cooling are known as convection, impingement and film cooling. U.S. Pat. Nos. 3,700,348 and 3,715,170, assigned to the assignee of the present invention, are excellent examples of advanced turbine air-cooling technology incorporating these basic air-cooling forms.
However, the benefits obtained from sophisticated air-cooling techniques are at least partially offset by the extraction of the necessary cooling air from the propulsive cycle. For example, probably the most popular turbine coolant today is air which is bled off of the compressor portion of the gas turbine engine and is routed to the hollow interior of the turbine blades. Typically, the work which has been done on this air by the compressor is partially lost to the cycle. Additionally, as the cooling air circulates throughout the turbine blade it picks up heat from the metallic blades or vanes. When his heated cooling air leaves the turbine blades, perhaps as a coolant film, this heat energy is lost to the cycle since the hot gases are normally mixed with the exhaust gases and ejected from an engine nozzle. It would be desirable, therefore, to have a cooling system wherein a medium other than compressor bleed air is used and wherein the heat extracted by the cooling medium is put back into the cycle in a useful and practical manner.
A partial solution to the foregoing problems has been the suggestion of closed-loop cooling systems for turbine blades which may or may not also incorporate the concept of regeneration or recuperation to recover lost thermal energy. One such cooling arrangement which has been proposed, for example, is that of U.S. Pat. No. 2,782,000. In that patent, a closed-system steam thermosiphon is used to cool the turbine blades, the thermosiphon principle being that by which a coolant is caused to circulate throughout the hollow bores of a turbine blade under the pumping of centrifugal force due to the difference in density between the heated coolant (steam) exiting the blade and the coolant (steam or water) entering it. Each blade is provided with its own thermosiphon which is associated with a cooler or heat exchanger which, in turn, is cooled by a second cooling medium such as water or air.
However, such closed-loop systems utilizing steam, sodium and potassium (for example) as coolants have a disadvantage in that if a leak develops in the coolant loop the cooling capability is lost. This disadvantage is not inherent in the air-cooled blades, which are typically of the open circuit type, wherein air extracted from the compressor is routed through the blades and discharged therefrom as a coolant film. Thus, it becomes desirable to provide a secondary means of cooling turbine blades which are normally cooled by the closed-loop cooling technique in the event that a leak should develop therein. Since compressor bleed air is normally available for this purpose, it is desirable to utilize this cooling air as the secondary coolant and to terminate the flow thereof except when absolutely necessary due to the power cycle implications.